1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to turbine airfoils with cooling circuits.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine is very efficient machine that converts the chemical energy of a burning fuel into mechanical energy. An industrial gas turbine (IGT) engine is used in power plants to drive an electric generator to produce electric power. An aero gas turbine engine is used to propel an aircraft. Both types of gas turbine engines benefit from increased efficiency. The efficiency of a gas turbine engine can be increased by increasing the high temperature gas flow that enters the turbine. It is a very important design feature to provide for the first stage stator vanes and rotor blades to have a high heat resistance as possible by using high temperature resistant materials in combination with internal and film cooling of the airfoils (vanes and blades).
Improved cooling of a turbine airfoil will allow for higher gas flow temperature and therefore increased engine efficiency. Maximizing the use of the cooling air will also increase the engine efficiency since less cooling air is required to pass through the turbine airfoils for cooling. Since the compressed cooling air used to cool the internal passages of the turbine airfoil is diverted from the compressor, using less cooling air to provide the same amount of cooling will also increase the engine efficiency.
In the cited prior art references, U.S. Pat. No. 6,379,118 B2 issued to Lutum et al on Apr. 30, 2002 entitled COOLED BLADE FOR A GAS TURBINE discloses in FIG. 3 of this patent a sinusoidal flow cooling passage extending along the wall of the blade to provide near wall cooling to the blade. The sinusoidal flow path is formed by an alternating series of ribs extending perpendicular from the wall surface in the flow direction. The sinusoidal flow path in the Lutum patent does not flow from the leading edge supply passage to the trailing edge region of the blade, nor does the sinusoidal flow passages or channels extend from the pressure side wall to the suction side wall as in the present invention.
Another prior art reference, U.S. Pat. No. 5,752,801 issued to Kennedy on May 19, 1998 entitled APPARATUS FOR COOLING A GAS TURBINE AIRFOIL AND METHOD OF MAKING SAME discloses a sinusoidal flow path extending along the trailing edge region of the blade, where the sinusoidal flow path extends from the pressure side wall to the suction side wall and out the trailing edge through exit holes. The sinusoidal flow passage in the Kennedy patent is one path extending along the spanwise length of the blade and occurs only in the trailing edge region of the blade.
Another prior art reference, U.S. Pat. No. 3,220,697 issued to Smuland et al on Nov. 30, 1965 entitled HOLLOW TURBINE OR COMPRESSOR VANE discloses a turbine airfoil with an internal cooling passage in FIG. 2 of this patent that follows a sinusoidal flow path from the outer shroud to the inner shroud, exiting out a hole in the inner shroud. The Smuland patent does not include a plurality of sinusoidal flow passages, nor does it show a sinusoidal flow passage extending along the blade or airfoil chordwise direction as does the present invention.
It is an object of the present invention to provide for a turbine airfoil, whether it be a stator vane or a rotor blade, with improved heat transfer coefficient from the hot metal to the cooling air passing through the cooling channels. It is also an object of the present invention to provide for a turbine airfoil that will provide the same amount of cooling for the airfoil but with less cooling air flow over the cited prior art references.